In modern commercial aircraft, high compression loads may be transitioned from an outer wing skin panel through a side-of-body (SOB) rib to an inner wing skin panel. On an aircraft with composite wings, composite stringers are commonly bonded to the composite wing skin panel. A large portion of the loads may be carried through stringers that are bonded to the wing skin panels. The inner and outer wing skin panels and stringers may be bolted to splice plates which are provided on the SOB or by some other suitable means to transfer loads between the wing skin panels. The offset of the splice plates relative to the centroid load path of the wing skin panels may induce a bending moment at the ends of the wing panels that is reacted as a pull-off load in the web of the stringer. The pull-off load may occur at the end of the stringer, where resistance to stress and strain concentrations may be needed. In metallic structure, this is less problematic due to the material system(s) isotropic properties with respect to in and out of plane loading. In composite structure, the out of plane loads act on the weaker laminate interface resulting in delamination of composites. In a typical composite skin/stringer design, stringers are co-bonded onto the skin panel. This configuration may create a situation where a large induced moment is applied along the skin/stringer bond line at the stringer free edge creating pull off forces. The intent of this trim is to reduce localized stress concentrations at the free edge of the stringer.
In structures in which the end of the composite wing stringer web which is attached to the SOB rib is solid from the stringer cap to the stringer base flange, the load path through the stringer end may impart undesired stress and strain to the cap and base flange of the stringer.
Accordingly, it would be advantageous to have an apparatus and method which takes into account one or more of the issues discussed above, as well as possibly other issues.